1. Field of the Invention
The present invention relates generally to a gas turbine engine, and more specifically to a turbine rotor blade with a showerhead film cooling hole arrangement.
2. Description of the Related Art Including Information Disclosed Under 37 CFR 1.97 and 1.98
A gas turbine engine includes a turbine section with a plurality of stages of stationary vanes and rotary blades to extract mechanical energy from a hot gas flow passing through the turbine. The gas turbine engine efficiency can be increased by providing for a higher temperature of the gas flow entering the turbine. The temperature entering the turbine is limited to the first stage vane and rotor blades ability to withstand the high temperature.
One method of allowing for higher temperatures than the material properties of the first stage vane and blades would allow is to provide for cooling air passages through the airfoils. Since the cooling air used to cool the airfoils is generally bled off from the compressor, it is also desirable to use a minimum amount of bleed off air in order to improve the efficiency of the engine. The compressor performs work to compress the bleed air for use in cooling the airfoils.
The hottest part of the airfoils is found on the leading edge. Complex designs have been proposed to provide the maximum amount of cooling for the leading edge while using the minimum amount of cooling air. One leading edge airfoil design is the showerhead arrangement. In the Prior Art, a blade leading edge showerhead comprises three rows of cooling holes as shown in FIG. 1. The showerhead arrangement 10 of the Prior Art includes a cooling air supply channel 11, a metering hole 13, a showerhead cavity 12, and a plurality of film cooling holes 14. The middle film row is positioned at the airfoil stagnation point which is where the highest heat load is found on the airfoil leading edge. The cooling hole labeled as 14 in FIG. 1 with the arrow indicates the cooling air flow is the stagnation point. The stagnation point is where the highest heat load appears on the airfoil leading edge. Film cooling holes for each row are at inline pattern and at staggered array relative to the adjacent film row as seen in FIG. 4. The showerhead cooling holes 14 are inclined at 20 to 35 degrees relative to the blade leading edge radial surface as shown in FIG. 3.
The Prior Art showerhead arrangement of FIGS. 1-4 suffers from the following problems. The heat load onto the blade leading edge region is in parallel to the film cooling hole array, and therefore reduces the cooling effectiveness. The portion of the film cooling holes within each film row is positioned behind each other as shown in FIG. 3 that reduces the effective frontal convective area and conduction distance for the oncoming heat load. Realistic minimum film hole spacing to diameter ratio n is approximately at 3.0. Below this ratio, zipper effect cracking may occur for the film row. This translates to maximum achievable film coverage for that particular film row to be 33% or 0.33 film effectiveness for each showerhead film row. Since the showerhead film holes are at radial orientation, film pattern discharge from the film hole is overlapped to each other. Little or no film is evident in-between film holes.
To allow for higher temperature exposure, a thin TBC (Thermal Barrier Coating) is used in the turbine airfoil leading edge cooling design to provide additional insulation for the airfoil for the reduction of heat load from the hot gas to the airfoil which reduces the airfoil metal temperature and thus reduces the cooling flow consumption and improves the turbine efficiency. As the turbine inlet temperature increases as turbines improve, the cooling flow demand for cooling the airfoil will increase and thus reduce the turbine efficiency. One alternative way for reducing the cooling air consumption while increasing the turbine inlet temperature for higher turbine efficiency is by using a thicker TBC on the cooled airfoil. Thus, the airfoil cooling design becomes more reliant on the endurance of the coating and thus the TBC becomes the prime design feature of the cooling design for the airfoil. A thicker TBC results in higher chances of spallation (when chips of the coating break away from the airfoil surface and leave exposed metal).